The present disclosure relates to coating processes and equipment for gas turbine engine components, and more particularly, to a method and masking assembly for selectively depositing a coating on a selected portion of a gas turbine engine component.
Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to burn hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases.
Gas turbine engine hot section components such as turbine blades and turbine vanes are subject to high thermal loads for prolonged time periods. Turbine section blades and vanes typically include an airfoil within hot combustion core gases from the upstream combustor section. Because of the high temperatures and corrosive effects of such hot combustion core gases a protective coating such as a thermal barrier coating provides insulation from the high temperatures and corrosive effects therefrom.
One method to coat only the flow path surfaces is to segregate the airfoil with a rectilinear mask that shields a generally rectilinear platform. An operator manually slides the generally rectilinear platform into the mask to surround the platform. The mask may form a gap with respect to a generally rectilinear platform due to component tolerances.
The mask is manufactured of a temperature resistant metallic alloy to survive the thermal barrier coating spray application. Oftentimes the gap may receive overspray from the thermal barrier coating spray application and may glue the mask to the component. Once the thermal barrier coating sets, the component is typically released by tapping the component from the mask. This tapping may undesirably chip the thermal barrier coating and thereby reduce protection of the component.